Ceramic matrix composite component cooling

ABSTRACT

Ceramic matrix composite (CMC) airfoils and methods for forming CMC airfoils are provided. In one embodiment, an airfoil is provided that includes opposite pressure and suction sides extending radially along a span and opposite leading and trailing edges extending radially along the span. The leading edge defines a forward end of the airfoil, and the trailing edge defines an aft end of the airfoil. A trailing edge portion is defined adjacent the trailing edge at the aft end, and a pocket is defined in and extends within the trailing edge portion. A heat pipe is received in the pocket. A method for forming an airfoil is provided that includes laying up a CMC material to form an airfoil preform assembly; processing the airfoil preform assembly; defining a pocket in a trailing edge portion of the airfoil; and inserting a heat pipe into the pocket.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of and claims priority to U.S.application Ser. No. 15/232,880, filed Aug. 10, 2016, the contents ofwhich are incorporated herein by reference.

FIELD OF THE INVENTION

The present subject matter relates generally to ceramic matrix compositecomponents and particularly to features for cooling ceramic matrixcomposite components of gas turbine engines. More particularly, thepresent subject matter relates to trailing edge cooling for ceramicmatrix component airfoils of gas turbine engines.

BACKGROUND OF THE INVENTION

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine generally includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gasesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

In general, turbine performance and efficiency may be improved byincreased combustion gas temperatures. Non-traditional high temperaturematerials, such as ceramic matrix composite (CMC) materials, are morecommonly being used for various components within gas turbine engines.For example, because CMC materials can withstand relatively extremetemperatures, there is particular interest in replacing componentswithin the flow path of the combustion gases with CMC materials.However, even though CMC components may withstand more extremetemperatures than typical components, CMC components still may requirecooling features or reduced exposure to the combustion gases to decreasea likelihood of negative impacts of increased combustion gastemperatures, e.g., material failures or the like.

More specifically, CMC airfoils for gas turbine engines typically have acavity for receipt of a cooling fluid located near a forward end of theairfoil, i.e., proximate a leading edge of the airfoil. Often, an aftend of the airfoil, i.e., proximate a trailing edge of the airfoil, doesnot have a cavity and is not near a cavity for receipt of a coolingfluid or other feature(s) for cooling the aft end, and thus, the aft endof the airfoil remains uncooled, which can produce a large temperaturegradient between the forward end and the aft end. A large temperaturegradient across the airfoil can increase the thermal stress or strain onthe airfoil, which can lead to material failures, reduced life of theairfoil, or other negative impacts on turbine performance.

Therefore, improved cooling features for gas turbine components, andspecifically CMC components for gas turbine engines, that overcome oneor more disadvantages of existing components would be desirable. Inparticular, an airfoil for a gas turbine engine having cooling featuresin a trailing edge portion of the airfoil would be beneficial. Moreover,a turbine nozzle for a gas turbine engine having cooling features in atrailing edge portion of an airfoil of the turbine nozzle that even outcooling of the airfoil would be desirable. Further, a CMC airfoil havingcooling features in a trailing edge portion of the airfoil would beuseful. Methods for forming a CMC airfoil of a gas turbine engine suchthat the airfoil has one or more cooling features in a trailing edgeportion of the airfoil also would be advantageous.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, an airfoil for agas turbine engine is provided. The airfoil includes opposite pressureand suction sides extending radially along a span and opposite leadingand trailing edges extending radially along the span. The pressure andsuction sides extend axially between the leading and trailing edges. Theleading edge defines a forward end of the airfoil, and the trailing edgedefines an aft end of the airfoil. The airfoil further includes atrailing edge portion defined adjacent the trailing edge at the aft endand a pocket defined in the trailing edge portion. The pocket extendswithin the trailing edge portion, and a heat pipe is received in thepocket.

In another exemplary embodiment of the present disclosure, a method forforming a ceramic matrix composite (CMC) airfoil for a gas turbineengine is provided. The method comprises laying up a CMC material toform an airfoil preform assembly. The airfoil preform assembly definesan airfoil shape having opposite pressure and suction sides extendingradially along a span and opposite leading and trailing edges extendingradially along the span. The pressure and suction sides extend axiallybetween the leading and trailing edges. The leading edge defines aforward end of the airfoil, and the trailing edge defines an aft end ofthe airfoil. The airfoil shape also has a trailing edge portion definedadjacent the trailing edge at the aft end. The method further comprisesprocessing the airfoil preform assembly to produce a green state CMCairfoil; defining a pocket in the trailing edge portion; and inserting aheat pipe into the pocket.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 provides a schematic cross-sectional view of an exemplary gasturbine engine according to various embodiments of the present subjectmatter.

FIG. 2 provides a perspective view of a turbine nozzle segment accordingto an exemplary embodiment of the present subject matter.

FIG. 3A provides a radial cross-section view of the turbine nozzlesegment of FIG. 2 according to an exemplary embodiment of the presentsubject matter.

FIG. 3B provides a portion of the cross-section of FIG. 3A, illustratingpockets defined in an airfoil for receiving heat pipes.

FIG. 3C provides a radial cross-section view of the turbine nozzlesegment of FIG. 2 according to another exemplary embodiment of thepresent subject matter.

FIG. 4A provides an axial cross-section view of the airfoil of theturbine nozzle segment of FIG. 2 according to an exemplary embodiment ofthe present subject matter.

FIG. 4B provides a perspective view of a heat pipe according to anexemplary embodiment of the present subject matter.

FIG. 5A provides a radial cross-section view of the turbine nozzlesegment of FIG. 2 according to another exemplary embodiment of thepresent subject matter.

FIG. 5B provides a portion of the cross-section of FIG. 5A, illustratingpockets defined in an airfoil for receiving heat pipes.

FIG. 5C provides a radial cross-section view of the turbine nozzlesegment of FIG. 2 according to another exemplary embodiment of thepresent subject matter.

FIG. 6A provides an axial cross-section view of the airfoil of theturbine nozzle segment of FIG. 2 according to another exemplaryembodiment of the present subject matter.

FIG. 6B provides a perspective view of a heat pipe according to anotherexemplary embodiment of the present subject matter.

FIG. 7A provides a radial cross-section view of the turbine nozzlesegment of FIG. 2 according to another exemplary embodiment of thepresent subject matter.

FIG. 7B provides a radial cross-section view of the turbine nozzlesegment of FIG. 2 according to another exemplary embodiment of thepresent subject matter.

FIG. 8 provides a chart illustrating a method for forming an airfoil ofa gas turbine engine according to an exemplary embodiment of the presentsubject matter.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first,” “second,” and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. In general, the turbofan 10includes a fan section 14 and a core turbine engine 16 disposeddownstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22.

For the depicted embodiment, fan section 14 includes a fan 38 having aplurality of fan blades 40 coupled to a disk 42 in a spaced apartmanner. As depicted, fan blades 40 extend outward from disk 42 generallyalong the radial direction R. Fan blades 40 and disk 42 are togetherrotatable about the longitudinal axis 12 by LP shaft 36.

Referring still to the exemplary embodiment of FIG. 1, disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated thatnacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersturbofan 10 through an associated inlet 60 of the nacelle 50 and/or fansection 14. As the volume of air 58 passes across fan blades 40, a firstportion of the air 58 as indicated by arrows 62 is directed or routedinto the bypass airflow passage 56 and a second portion of the air 58 asindicated by arrows 64 is directed or routed into the LP compressor 22.The ratio between the first portion of air 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

In some embodiments, components of turbofan engine 10, particularlycomponents within hot gas path 78, such as components of the combustionand/or turbine sections, may comprise a ceramic matrix composite (CMC)material, which is a non-metallic material having high temperaturecapability. Exemplary CMC materials utilized for such components mayinclude silicon carbide (SiC), silicon nitride, or alumina matrixmaterials and combinations thereof. Ceramic fibers may be embeddedwithin the matrix, such as oxidation stable reinforcing fibers includingmonofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6),as well as roving and yarn including silicon carbide (e.g., NipponCarbon's NICALON®, Ube Industries' TYRANNO®, and Dow Corning'sSYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480), and choppedwhiskers and fibers (e.g., Nextel's 440 and SAFFIL®), and optionallyceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinationsthereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica,talc, kyanite, and montmorillonite). For example, in certainembodiments, bundles of the fibers, which may include a ceramicrefractory material coating, are formed as a reinforced tape, such as aunidirectional reinforced tape. A plurality of the tapes may be laid uptogether (e.g., as plies) to form a preform component. The bundles offibers may be impregnated with a slurry composition prior to forming thepreform or after formation of the preform. The preform may then undergothermal processing, such as a cure or burn-out to yield a high charresidue in the preform, and subsequent chemical processing, such asmelt-infiltration with silicon, to arrive at a component formed of a CMCmaterial having a desired chemical composition. In other embodiments,the CMC material may be formed as, e.g., a carbon fiber cloth ratherthan as a tape.

CMC materials may be used for various components of the engine, forexample, airfoils, shrouds, and/or other components in the compressor,turbine, and/or fan regions. As a particular example, turbine nozzles,comprising stator vanes extending between inner and outer bands, directthe hot combustion gases in a manner to maximize extraction at theadjacent downstream turbine blades. As such, CMC materials are desirablefor use in forming turbine nozzles exposed to the high temperatures ofthe hot combustion gases. Of course, other components of turbine engine10 also may be formed from CMC materials.

Referring now to FIG. 2, a perspective view is provided of a turbinenozzle segment 100 according to an exemplary embodiment of the presentsubject matter. A turbine stator is formed by a plurality of turbinenozzle segments 100 that are abutted at circumferential ends, e.g., endsor sides spaced apart along a circumferential direction C, to form acomplete ring about centerline 12. Each nozzle segment 100 may comprisean inner band 102 and an outer band 104 with one or more vanes 106extending from inner band 102 to outer band 104. In some embodiments,vanes 106 may be vanes 68 of HP turbine 28 or vanes 72 of LP turbine 30described above. Each stator vane 106 includes an airfoil 108 having aconcave pressure side 110 (FIGS. 4A, 6A) opposite a convex suction side112. Opposite pressure and suction sides 110, 112 of each airfoil 108extend radially along a span S (FIGS. 3A, 5A) from a vane root at innerband 102 to a vane tip at outer band 104. Moreover, pressure and suctionsides 110, 112 of airfoil 108 extend axially between a leading edge 114and an opposite trailing edge 116. Leading edge 114 defines a forwardend of airfoil 108 (labeled FWD in the Figures), and trailing edge 116defines an aft end of airfoil 108 (labeled AFT in the Figures). Pressureand suction sides 110, 112 of airfoil 108 define an outer surface 118 ofthe airfoil. Additionally, airfoil 108 may define one or more cavities115 (e.g., as shown in FIG. 3A) adjacent leading edge 114 for receivinga flow of cooling fluid, e.g., a flow of pressurized air diverted fromHP compressor 24. As such, each cavity 115 may provide cooling to theportion of airfoil 108 adjacent the leading edge 114. In an exemplaryembodiment, airfoil 108 is formed from a CMC material, e.g., asdescribed in greater detail below, such that airfoil 108 is a CMCairfoil. In other embodiments, airfoil 108 may be formed from anotherappropriate material.

FIG. 3A provides a radial cross-section view of turbine nozzle segment100 according to an exemplary embodiment of the present subject matter.FIG. 3B provides a cross-sectional view of a portion of airfoil 108 ofturbine nozzle segment 100 according to an exemplary embodiment of thepresent subject matter. As shown in FIGS. 3A and 3B, airfoil 108includes a trailing edge portion 120 that is defined adjacent thetrailing edge 116 at the aft end of airfoil 108. Referring particularlyto FIG. 3B, a plurality of pockets 122 are defined in the trailing edgeportion 120. Each pocket 122 extends axially within the trailing edgeportion 120. More specifically, in the illustrated embodiment, eachpocket 122 extends axially from an aft portion of cavity 115 toward thetrailing edge 116 of airfoil 108 such that a forward end 122F of eachpocket 122 is defined at cavity 115 and an aft end 122A of each pocket122 is defined aft of the cavity 115 toward trailing edge 116. Further,the plurality of pockets 122 are radially spaced apart from one another,i.e., the plurality of pockets 122 are individually defined along theradial direction R.

As depicted in FIG. 3A, a heat pipe 124 is received in each pocket ofthe plurality of pockets 122. Each heat pipe 124 includes a body havingan evaporator portion 124E and a condenser portion 124C. Heat isabsorbed or transferred from the trailing edge portion 120 intoevaporator portion 124E of heat pipe 124, particularly a vaporizableliquid in heat pipe 124. The heat is then dissipated through condenserportion 124C into its environment, e.g., into cavity 115 and/or into arelatively cooler segment of airfoil 108 than the segment near trailingedge 116. That is, in embodiments such as the embodiment depicted inFIG. 3A, the condenser portion 124C of each heat pipe may be positionedat the cavity 115 to dispel heat from the trailing edge portion 120 ofairfoil 108 into the cavity 115. In some embodiments, a fluid, such ascooling fluid F within cavity 115, may flow over condenser portion 124Cand thereby dissipate the heat more quickly. As a particular example,FIG. 3C depicts an embodiment in which the condenser portion 124C ofeach heat pipe 124 extends into cavity 115. In embodiments such as theembodiment of FIG. 3C, the cooling fluid F may more readily flow overthe condenser portion 124C, which can help more quickly dissipate heatabsorbed at the evaporator portion 124E of the heat pipe. In otherembodiments, the condenser portion 124C may extend to, but not into,cavity 115 such that the cooling fluid F flows over an end of condenserportion 124C positioned at cavity 115, as illustrated in FIG. 3A.

Each heat pipe 124 uses a liquid that evaporates by absorbing the heatfrom a hot end, i.e., evaporator portion 124E. The vapor generated thentravels through a channel formed within the body of the heat pipe 124that extends from the evaporator portion 124E to the condenser portion124C, and the vapor condenses at the cold end, i.e., condenser portion124C, thereby transferring heat to the cold end. A capillary structureor wick that extends from one end of the heat pipe to the other issaturated with a volatile or working fluid to transport the condensedliquid back to the hot end by capillary action, thus completing thecircuit.

A working fluid and a body material of heat pipe 124 may be selectedbased on operating conditions of turbofan engine 10 and, moreparticularly, core turbine engine 16. In one exemplary embodiment, heatpipe 124 may experience operating temperatures between about 1000° C.and 1100° C. In such an embodiment, suitable working fluids includelithium (Li) and sodium (Na), although other working fluids also may beappropriate. Further, the body of heat pipe 124 may be made from aniobium-zirconium (Nb—Zr) alloy such as Nb-1% Zr, or from a sinteredniobium-zirconium (Nb—Zr) alloy such as Nb-1% Zr, e.g., if the workingfluid is sodium. The axial heat flux of each heat pipe 124 constructedusing these working fluids and body materials may be about 200 W/cm².Other heat flux values may be possible using other working fluids orbody materials.

FIG. 4A provides an axial cross-section view of airfoil 108 according toan exemplary embodiment of the present subject matter. FIG. 4B providesa perspective view of a heat pipe 124 according to an exemplaryembodiment of the present subject matter. As shown in FIGS. 4A and 4B,each heat pipe 124 may have a generally trapezoidal shape, e.g., suchthat the heat pipe tapers to fit within airfoil 108 as the airfoiltapers along its width in the trailing edge portion of the airfoil. Inother embodiments, however, heat pipe 124 may have any appropriateshape, e.g., heat pipe 124 may be generally cylindrical or polyhedral inshape.

Turning to FIGS. 5A and 5B, in another embodiment of airfoil 108 andturbine nozzle section 100, heat pipes 124 may extend radially ratherthan axially within trailing edge portion 120 of airfoil 108. Asillustrated in FIG. 5B, one or more pockets 122 may be defined in thetrailing edge portion 120 of airfoil 108 such that each pocket 122extends radially within the trailing edge portion 120. Moreparticularly, in the depicted embodiment of FIG. 5B, each pocket 122extends radially within trailing edge portion 120 adjacent an aftportion of cavity 115 such that an inner end 122I of each pocket 122 isdefined closer to the airfoil root than an outer end 122O, which isdefined closer to the airfoil tip than the inner end 122I of therespective pocket. Further, each pocket 122 is radially spaced apartfrom another pocket 122 defined within the airfoil 108, i.e., eachpocket 122 is individually defined along the radial direction R.

As depicted in FIG. 5A, one or more pockets 122 may be defined inairfoil 108, and a heat pipe 124 may be received in each pocket 122. Asdescribed above with respect to FIG. 3A, each heat pipe 124 includes abody having an evaporator portion 124E and a condenser portion 124C.Heat is absorbed or transferred from the trailing edge portion 120 intoevaporator portion 124E of heat pipe 124, particularly a vaporizableliquid in heat pipe 124. The heat is then dissipated through condenserportion 124C into its environment, e.g., into a relatively coolersegment of airfoil 108 such as an inner portion of the airfoil, an outerportion of the airfoil, or a midsection of the airfoil. For example, asillustrated in FIG. 5A, the condenser portion 124C of a radially outerheat pipe 124 may be positioned near the radially outermost portion ofairfoil 108 to dispel heat from the trailing edge portion 120 of airfoil108 into a radially outer portion of the airfoil and/or into outer band104. Further, the condenser portion 124C of a radially inner heat pipe124 may be positioned near the radially innermost portion of airfoil 108to dispel heat from the trailing edge portion 120 of airfoil 108 into aradially inner portion of the airfoil and/or into inner band 102. Assuch, heat pipes 124 may dissipate heat from a relatively warmermidsection of the airfoil 108 to relatively cooler inner and outerportions of the airfoil. In other embodiments, the condenser portions124C of the radially inner and outer heat pipes 124 may be positionednear the midsection of the airfoil 108 to dissipate heat from relativelywarmer inner and outer portions of the airfoil to a relatively coolermidsection of the airfoil.

As another example, as illustrated in FIG. 5C, the condenser portion124C of each heat pipe 124 may extend into a cavity defined by, e.g.,inner band 102 or outer band 104. In such embodiments, a cooling fluid Fwithin the cavity may flow over the condenser portion 124C, which canhelp more quickly dissipate heat absorbed at the evaporator portion 124Eof the heat pipe. In other embodiments, disposing each heat pipe 124such that condenser portion 124C extends into a cavity away from thetrailing edge portion 120 of airfoil 108 may provide a better heat sinkfor more efficiently dissipating heat than disposing the condenserportion 124C within the airfoil 108. Accordingly, the heat pipes 124 maybe received in pockets 122 in any appropriate orientation for balancingthe thermal gradient experienced by a given airfoil.

As described above, each heat pipe 124 uses a liquid that evaporates byabsorbing the heat from the evaporator portion 124E, i.e., the hot end,and thereby generating vapor that then travels through a channel formedwithin the body of the heat pipe 124 and the vapor condenses at thecondenser portion 124C, i.e., the cold end. As such, each heat pipetransfers heat from the hot end to the cold end. A capillary structureor wick that extends from one end of the heat pipe to the other issaturated with a volatile or working fluid to transport the condensedliquid back to the hot end by capillary action, thus completing thecircuit. The working fluid and body material of heat pipe 124 may beselected based on operating conditions of turbofan engine 10 and, moreparticularly, core turbine engine 16. As previously described, suitableworking fluids include lithium (Li) and sodium (Na) and suitable heatpipe body materials include a niobium-zirconium (Nb—Zr) alloy such asNb-1% Zr or a sintered niobium-zirconium (Nb—Zr) alloy such as Nb-1% Zr,although other working fluids and body materials also may beappropriate. The radial heat flux of each heat pipe 124 constructedusing these working fluids and body materials may be about 200 W/cm².Other heat flux values may be possible using other working fluids orbody materials.

FIG. 6A provides an axial cross-section view of the airfoil 108 of FIG.5A, and FIG. 6B provides a perspective view of a heat pipe 124 accordingto an exemplary embodiment of the present subject matter. As shown inFIGS. 6A and 6B, each heat pipe 124 may have a generally cylindricalshape, where a length of each cylinder extends along the radialdirection R. In other embodiments, heat pipe 124 may have anyappropriate shape, e.g., heat pipe 124 may be generally trapezoidal asshown in FIGS. 4A and 4B, or heat pipe 124 may be polyhedral in shape.

Turning now to FIGS. 7A and 7B, heat pipes 124 may be disposed withintrailing edge portion 120 in other configurations as well. Referringparticularly to FIG. 7A, some or all of heat pipes 124 may be disposedat an angle with respect to the axial direction A and radial directionR. As depicted in the embodiment of FIG. 7A, one heat pipe 124 extendsgenerally along or parallel to the axial direction A while the otherheat pipes 124 within trailing edge portion 120 extend at an angle withrespect to the axial direction A and at an angle with respect to theradial direction R. In other embodiments, one or more heat pipes 124 mayextend generally along or parallel to the axial direction A and/or theradial direction R while one or more other heat pipes 124 are disposedat an angle with respect to the axial and radial directions A, R. Instill other embodiments, referring to FIG. 7B, a plurality of heat pipes124 may be disposed within trailing edge portion 120 such that a portionof the heat pipes 124 extend generally along or parallel to the axialdirection A and the remaining portion of the heat pipes 124 extendgenerally along or parallel to the radial direction R. In the particularembodiment of FIG. 7B, the radially extending heat pipes 124 aredisposed in a radially outer portion of airfoil 108, with the condenserportion 124C of each heat pipe 124 disposed within a cavity defined byouter band 104. However, in other embodiments, the condenser portion124C may not extend past outer band 104 into the cavity, or the radiallyextending heat pipes 124 may be disposed within a radially inner portionof airfoil 108 with the condenser portion 124C of each heat pipeextending toward or into a cavity defined by the inner band 102. Otherorientations or combinations of orientations of heat pipes 124 withinairfoil 108 may be used as well.

In various embodiments of airfoil 108, such as the embodimentsillustrated in FIGS. 3A, 3C, 5A, 5C, 7A, and 7B, airfoil 108 includesmore than one heat pipe 124. In such embodiments, each heat pipe 124 mayhave the same shape and size or the plurality of heat pipes 124 may varyin shape and/or size, i.e., each heat pipe need not be the same shape orsize. The shape, size, number, position, and orientation of heat pipes124 may be optimized for each airfoil. For example, the shape, size,number, position, and orientation of heat pipes 124 included within agiven airfoil 108 may depend on the relative size of the airfoil, suchthat, e.g., airfoils of first stage turbine nozzles may have a differentshape, size, number, position, and/or orientation of heat pipes 124 thanairfoils of second stage turbine nozzles. Moreover, the size, shape,number, position, and/or orientation of heat pipes 124 may depend on thedesired cooling effects achieved by the heat pipes. For example,increasing the size and/or number of heat pipes 124 within a givenairfoil 108 may enhance the cooling of the trailing edge portion 120 ofthe airfoil 108. However, defining too many pockets 122 within theairfoil 108 can be detrimental to the strength of the material formingthe airfoil. Therefore, an optimal number, shape, size, position, andorientation of heat pipes 124, as well as the corresponding pockets 122that receive the heat pipes, provides beneficial cooling without overlyweakening the airfoil material or otherwise negatively impacting engineperformance.

Heat pipes 124, whether extending axially, radially, or otherwise withinairfoil 108, provide cooling to trailing edge portion 120, e.g., byproviding increased thermal gradient control to reduce thermal stressesin airfoil 108. That is, heat pipes 124 located in trailing edge portion120 can help even out temperature gradients in airfoil 108 to renderairfoil 108 more isothermal than airfoil 108 without heat pipes 124.Particularly in airfoil 108 having cavity 115 that receives coolingfluid adjacent the leading edge 114 of the airfoil, balancing thethermal gradients of airfoil 108 along the axial direction A by alsoproviding cooling adjacent the trailing edge 116 via heat pipes 124 mayhelp improve the life of airfoil 108, as well as the performance of gasturbine engine 10. As described above, heat pipes 124 also may helpbalance the thermal gradients of airfoil 108 along the radial directionR. Further, as shown, e.g., in FIGS. 3A and 5A, the heat pipe(s) 124provided in airfoil 108 may be self-contained within the airfoil, i.e.,the one or more heat pipes need not protrude into another component ofthe turbofan engine 10 such as, e.g., within a stream of air from fansection 14.

Thus, a method of cooling an airfoil 108 of, e.g., a turbine nozzlesegment 100, includes providing heat pipe(s) 124 within a trailing edgeportion 120 of the airfoil 108. The method may include definingpocket(s) 122 in the trailing edge portion 120 and further may includeinserting a heat pipe 124 into each pocket 122. The method also maycomprise orienting the heat pipe(s) 124 such that an evaporator portion124E of each heat pipe is adjacent or within a relatively warm portionof the trailing edge portion 120 and a condenser portion 124C of eachheat pipe is adjacent or within a relatively cool portion of thetrailing edge portion 120. As such, each heat pipe 124 may dissipateheat from the relatively warm portion of the trailing edge portion 120of the airfoil to the relatively cool portion of the trailing edgeportion 120.

Turning back to FIGS. 4A and 6A, airfoil 108 may be a CMC component ofengine 10. In some embodiments, inner and outer bands 102, 104 also maybe made from a CMC material such that each turbine nozzle segment 100 isa CMC component of engine 10. In the embodiments depicted in FIGS. 4Aand 6A, pressure and suction sides 110, 112 of airfoil 108 are definedby a first plurality of CMC plies 150, which also may be referred to asairfoil plies 150. Airfoil 108 further comprises a second plurality ofCMC plies 152 defining cavity 115 within airfoil 108; the secondplurality of plies 152 also may be referred to as cavity plies 152. Eachof the plurality of airfoil plies 150 extends from pressure side 110 tosuction side 112 of airfoil 108. Cavity plies 152 define cavity 115between pressure and suction sides 110, 112, i.e., within airfoil 108.One or more filler packs 154 are positioned between airfoil plies 150and cavity plies 152 within trailing edge portion 120 of airfoil 108. Itwill be appreciated that filler pack(s) 154 also may be positionedbetween airfoil and cavity plies 150, 152 within other portions ofairfoil 108. In other embodiments, filler pack(s) 154 may be omitted,and airfoil 108 and its features may be defined by airfoil plies 150 ora combination of airfoil plies 150 and cavity plies 152.

Preferably, but not necessarily, airfoil and cavity plies 150, 152contain continuous CMC fibers along their lengths. Continuous fiber CMCplies can help avoid relying on the interlaminar capability of theairfoil material to resist stresses on the airfoil. The continuousfibers may be maintained, e.g., by wrapping each airfoil ply 150 fromone of pressure and suction sides 110, 112 to the other of pressure andsuction sides 110, 112 around one or both of leading and trailing edges114, 116. Cavity plies 152 may be wrapped around a mandrel or otherappropriate support to help maintain continuous fibers in plies 152.

It should be appreciated that, in general, filler packs 154 may beformed from any suitable material and/or by using any suitable process.For example, in several embodiments, each filler pack 154 may be formedfrom a suitable fiber-reinforced composite material, such as a carbon orglass fiber-reinforced composite material. For instance, one or morefabric plies may be wrapped in a suitable manner to form one or morefiller packs 154 defining the desired shape of an interior of airfoil108, such as by shaping suitable ply packs to form each filler pack 154.In another embodiment, discontinuous materials, such as short or choppedfibers, particulates, platelets, whiskers, etc., may be dispersedthroughout a suitable matrix material and used to form each filler pack154.

Additionally, it should be appreciated that, in several embodiments,each filler pack 154 may correspond to a pre-fabricated component. Insuch embodiments, the filler pack(s) 154 may be installed within theinterior of airfoil 108 during or following manufacturing of the nozzlesegment 100. Alternatively, each filler pack 154 may be assembled orotherwise formed within airfoil 108. For instance, when filler pack 154is formed from one or more fabric plies, the plies may be laid up withinairfoil 108 together with the plies being used to create the airfoilstructure, e.g., airfoil plies 150 and cavity plies 152.

Various methods, techniques, and/or processes may be used to formpockets 122 in airfoil 108. For example, in embodiments such as theexemplary embodiment of FIG. 4A, a portion of each pocket 122 is definedin cavity plies 152, and the portion of each pocket 122 may be definedin the cavity plies 152 by cutting each individual cavity ply 152 beforeplies 152 are laid up as part of forming airfoil 108. Plies 152 may becut, e.g., using a precision Gerber cutter by Gerber Technology ofTolland, Conn. In other embodiments, another type of cutter or othermeans may be used to form cut-outs in cavity plies 152 to define atleast a portion of each pocket 122. Alternatively or additionally, atleast a portion of each pocket 122 may be defined using electricaldischarge machining (EDM), e.g., EDM drilling, or laser machining,precision machining, or other suitable machining technique or process.For example, using an EDM drilling process, each pocket 122 may bedefined through cavity plies 152 and/or in one or more filler packs 154.

In still other embodiments, at least a portion of pockets 122 may beformed using one or more fugitive material inserts. As an example, aninsert made from a fugitive material may be in a desired form (e.g.,shape, size, etc.) to define an axially or radially extending pocket122. The fugitive material insert is positioned within the layup asairfoil plies 150, cavity plies 152, and/or filler pack(s) 154 are laidup to form airfoil 108. In some embodiments, the insert may be formed ofSiC fibers in a silica carbide matrix. The insert may be one of variousforms, such as a tape cast, a preformed silicon dioxide tube, or a rapidprototype polymer coated with boron nitride, and the insert may beformed in various manners, e.g., sprayed, screen printed, or injectionmolded. It may be desirable that the fugitive material insert be a lowmelting metal or alloy that may melt during a burnout pyrolysisoperation or melt infiltration of a CMC layup preform, to thereby leavea void in the preform. In alternative embodiments, the fugitive materialinsert may be formed of a high temperature material that will not meltduring the burnout pyrolysis operation. Such high temperature materialinserts may be placed into the CMC during layup as a flexible tapefilled with powders of the high temperature materials. Alternately, allof the high temperature material inserts may be placed into the CMCduring layup as a dense, flexible wire or an inflexible rod or tube.Such high temperature materials, after the CMC component is meltinfiltrated, may require a subsequent air heat treatment to oxidize thehigh temperature material, a vacuum heat treatment, an inert gas heattreatment, an acid treatment, a base treatment, combinations thereof, oralternating combinations thereof, to remove the fugitive material. Thus,the fugitive material may be removed by melting, dissolution,sublimation, evaporation, or the like, and various materials aresuitable for use as the insert, such as materials that exhibitnon-wetting of the CMC preform, low or no reactivity with theconstituents of the CMC preform, and/or are completely fusible anddrainable at a temperature of a thermal treatment performed on the CMCpreform. In one example embodiment, fugitive material inserts fordefining pockets 122 are formed of fused silicon dioxide (SiO₂) in atubular shape, i.e., as quartz tubes, which may be positioned in anarray within trailing edge portion 120 of a layup of plies 150, 152,and/or filler pack(s) 154 for forming airfoil 108. Following a meltinfiltration process, the fused silicon dioxide is reduced to SiO andleaves the CMC component with voids forming pockets 122, into which heatpipes 124 may be inserted.

FIG. 8 provides a chart illustrating an exemplary method 800 forfabricating airfoil 108. As represented at 802 in FIG. 8, plies 150, 152and filler pack(s) 154 are laid up in the form of airfoil 108, i.e.,laid up in a desired shape to produce an airfoil preform assembly. Thelayup step or portion of the process thus may be referred to as thelayup preforming step and generally may comprise layering multiple pliesor structures, such as plies pre-impregnated with matrix material(prepreg plies), prepreg tapes, or the like, to form a desired shape ofthe resultant CMC component, e.g., airfoil 108. The layers are stackedto form a layup or preform, which is a precursor to the CMC component.

In some embodiments, multiple layups or preforms may be laid up togetherto form the airfoil preform assembly. More particularly, the layupportion 802 of method 800 may include laying up multiple preforms,filler packs, and/or plies to form the airfoil preform assembly. In anexemplary embodiment, the layup portion 802 may include forming a cavitypreform and one or more filler pack preforms, which are laid up withairfoil plies 150 as shown at 802 a, 802 b, and 802 c in FIG. 7 toproduce the airfoil preform assembly. To form the cavity preform, cavityplies 152 may be laid up, e.g., in or on a layup tool, mandrel, or mold,to generally define the shape of cavity 115 of airfoil 108. The cavitypreform may be compacted at atmosphere, i.e., at room temperature, andthen processed in an autoclave. The autoclave processing may beperformed at a reduced temperature compared to a standard autoclavecycle such that the cavity preform retains some flexibility andmalleability after autoclaving. Such flexibility and malleability mayhelp in defining voids (such as a portion of pockets 122) in the cavitypreform and/or laying up the cavity preform with other preforms and/orplies to produce the airfoil preform assembly. In some embodiments, thecompaction and/or autoclaving steps may be omitted, i.e., the compactionand autoclaving are optional, such that defining the cavity preformcomprises laying up cavity plies 152 without additional processing.

The layup preforming shown at 802 in FIG. 7 further may include formingone or more filler pack preforms. For example, filler pack material 154may be laid up, e.g., in or on a layup tool, mandrel, or mold, to defineone or more filler pack preforms. Next, each filler pack preform 154Pmay be compacted, e.g., at atmosphere as described above with respect tothe cavity preform. Then, the filler pack preform(s) may be processed inan autoclave, e.g., at a reduced temperature relative to a standardautoclave cycle such that filler pack preform(s) retain some flexibilityand malleability after autoclaving. The flexibility and malleability mayhelp in defining voids in the filler pack preform(s) and/or in laying upthe filler pack preform(s) with other preforms and/or plies to form theairfoil preform assembly. More particularly, after autoclaving at areduced temperature, the filler pack preform(s) are in a green state andretain some flexibility and malleability that can assist in furthermanipulation of the preform. For example, the voids forming pockets 122in trailing edge portion 120 of the resultant airfoil 108 may bemachined in the green state filler pack preform(s), and the malleabilityof green state preform may help in machining the pockets 122. In someembodiments, pockets 122 may be formed by machining two green statefiller pack preforms, such that one preform defines a first half of eachpocket 122 and the second preform defines a second half of each pocket122. As previously described, the pocket(s) 122 may be formed in fillerpack preform(s) using one or more of laser drilling or machining, EDM,cutting, precision machining, or other machining methods. In otherembodiments, one or more of the pockets 122 may be formed using fugitivematerial inserts that are laid up with the filler pack preform.

Laying up the CMC material to produce the airfoil preform assembly alsomay include laying up airfoil plies 150 with the cavity preform orcavity plies 152 and/or with the filler pack preform(s) or fillerpack(s) 154. It will be appreciated that airfoil plies 150 generallydefine the shape of pressure and suction sides 110, 112 of the resultantairfoil 108. Accordingly, at the layup preforming portion 802 ofexemplary method 800, a cavity preform or cavity plies 152, filler packpreform(s), filler pack(s) 154, and/or airfoil plies 150 may be laid uptogether to form an airfoil preform assembly. As previously describe, insome embodiments, one or more fugitive material inserts may bepositioned within the layers of the airfoil preform assembly to form oneor more of pockets 122 within airfoil 108.

Next, the airfoil preform assembly is processed as shown at 804 in FIG.7. For example, the airfoil preform assembly may be processed in anautoclave using a standard autoclave process. As such, the airfoilpreform assembly may be processed at a higher temperature than thefiller pack preform and the cavity preform described above. Afterprocessing, the airfoil preform assembly is in a green state. If pockets122, or a portion of pockets 122, have not been formed in cavity plies152, the cavity preform, filler pack(s) 154, and/or the filler packpreform(s) as described above, the pockets 122 (or the portion ofpockets 122 that remains to be defined) may be defined in the greenstate airfoil preform assembly. In various embodiments, defining pockets122 in the airfoil preform assembly may comprise using one or more oflaser drilling or machining, EDM drilling, cutting, or other machiningmethods to define the pockets.

Then, as shown at 806 in FIG. 7, the airfoil preform assembly mayundergo a burn-out cycle, i.e., a burn-out cycle may be performed. In anexample burn-out cycle, any mandrel-forming materials, as well ascertain fugitive materials or other meltable materials such asadditional binders in the CMC plies, are melted to remove suchmaterials. During burn-out, the CMC airfoil preform assembly may bepositioned to allow the melted materials to run out of the preform andthus remove the materials from the preform.

Next, as illustrated at 808, the CMC airfoil preform assembly may besubjected to one or more post-processing cycles for densification of thepreform assembly. Densification may be performed using any knowndensification technique including, but not limited to, Silcomp, meltinfiltration (MI), chemical vapor infiltration (CVI), polymerinfiltration and pyrolysis (PIP), and oxide/oxide processes.Densification can be conducted in a vacuum furnace having an establishedatmosphere at temperatures above 1200° C. to allow silicon or othermaterials to melt-infiltrate into the preform component.

Additionally or alternatively, after burn-out and densifying steps 806,808, the airfoil 108 may be manipulated mechanically or chemically asshown at 810 in FIG. 7 to remove any remaining fugitive materialinserted into the preformed shape during the layup preforming portion ofmethod 800. In some cases, the heat treatment may be used to oxidize theinsert to an oxide that may be melted or dissolved in an acid or base.In other embodiments, the insert may be directly dissolved in acid orbase, or otherwise chemically dissolved. In further embodiments, theinsert may be sublimed or evaporated in a vacuum heat treatment. Instill other embodiments, the insert may be oxidized and subsequentlysublimed or evaporated in a vacuum heat treatment. Mechanical methodsmay be used to mechanically remove the insert, and such mechanicalmethods may or may not be used with any of the previously describedmethods. Various chemical methods may be utilized as well. Of course, inembodiments in which pockets 122 are not formed using a fugitivematerial, the process of removing the fugitive material as illustratedat 810 in FIG. 7 may be omitted.

After any remaining fugitive material is removed, airfoil 108 may befinished as shown at 812. Finishing the airfoil 108 may include finishmachining the airfoil and/or applying an environmental barrier coating(EBC) to the airfoil. Other processes or steps also may be performed tofinish airfoil 108 and prepare the airfoil for use in turbofan engine10.

As described above, a heat pipe 124 may be inserted into each pocket 122defined in the CMC airfoil 108 such that a heat pipe 124 is received ineach pocket 122. The heat pipe(s) 124 may be inserted into pocket(s) 122at any appropriate point within method 800. For example, the heatpipe(s) 124 may be inserted after the CMC airfoil 108 is finished asillustrated at 812 in FIG. 7. In other embodiments, the heat pipe(s) 124may be inserted after any remaining fugitive material is removed at 810but before the airfoil 108 is finished at 812. Other points withinmethod 800 also may be appropriate for inserting heat pipe(s) 124.

Method 800 is provided by way of example only; it will be appreciatedthat some steps or portions of method 800 may be performed in anotherorder or may be omitted or repeated as needed. Additionally, othermethods of fabricating or forming airfoil 108 may be used as well. Inparticular, other processing cycles, e.g., utilizing other known methodsor techniques for compacting CMC plies, may be used. Further, when innerand outer bands 102, 104 are formed from CMC materials, similar methodsas described above with respect to method 800 may be used to form theinner band 102 and/or the outer band 104. Moreover, after inner band102, outer band 104, and airfoil 108 are fabricated from a suitablematerial, the turbine nozzle segment 100 is assembled such that airfoil108 extends from inner band 102 to outer band 104. In such embodiments,as well as in other appropriate embodiments, heat pipe(s) 124 may beinserted in pocket(s) 122 of airfoil 108 before the airfoil is assembledwith the inner and outer bands 102, 104. In appropriate embodiments,turbine nozzle segment 100 may be formed from a CMC material such thatthe inner band 102, outer band 104, and airfoil 108 are a single,unitary component. This written description uses examples to disclosethe invention, including the best mode, and also to enable any personskilled in the art to practice the invention, including making and usingany devices or systems and performing any incorporated methods. Thepatentable scope of the invention is defined by the claims and mayinclude other examples that occur to those skilled in the art. Suchother examples are intended to be within the scope of the claims if theyinclude structural elements that do not differ from the literal languageof the claims or if they include equivalent structural elements withinsubstantial differences from the literal language of the claims.

What is claimed is:
 1. An airfoil for a gas turbine engine, comprising:opposite pressure and suction sides extending radially along a span;opposite leading and trailing edges extending radially along the span,the pressure and suction sides extending axially between the leading andtrailing edges, the leading edge defining a forward end of the airfoil,the trailing edge defining an aft end of the airfoil; a trailing edgeportion defined adjacent the trailing edge at the aft end; at least twopockets defined in the trailing edge portion, the at least two pocketsextending within the trailing edge portion, the at least two pocketsradially spaced apart from one another; and a heat pipe received in eachpocket.
 2. The airfoil of claim 1, wherein at least one pocket of the atleast two pockets extends at an angle with respect to an axialdirection.
 3. The airfoil of claim 2, wherein the airfoil definesadjacent the leading edge a cavity for receipt of a flow of coolingfluid, and wherein a condenser portion of the heat pipe is positionednear the cavity.
 4. The airfoil of claim 1, wherein at least one pocketof the at least two pockets extends radially within the trailing edgeportion, and wherein at least one pocket of the at least two pocketsextends axially within the trailing edge portion.
 5. The airfoil ofclaim 4, wherein the airfoil defines adjacent the leading edge a cavityfor receipt of a flow of cooling fluid, and wherein a condenser portionof the at least one heat pipe extending axially within the trailing edgeportion is positioned near the cavity.
 6. The airfoil of claim 4,wherein the airfoil defines adjacent the leading edge a cavity forreceipt of a flow of cooling fluid, and wherein a condenser portion ofthe at least one heat pipe extending axially within the trailing edgeportion is positioned near an outer band of the airfoil.
 7. The airfoilof claim 4, wherein the airfoil defines adjacent the leading edge acavity for receipt of a flow of cooling fluid, and wherein a condenserportion of the at least one heat pipe extending axially within thetrailing edge portion is positioned near an inner band of the airfoil.8. The airfoil of claim 1, wherein each heat pipe comprises a condenserportion, and wherein the condenser portion of at least one heat pipe ofthe at least two heat pipes is positioned near the radially outermostportion of the airfoil.
 9. The airfoil of claim 1, wherein each heatpipe comprises a condenser portion, and wherein the condenser portion ofat least one heat pipe of the at least two heat pipes is positioned nearthe radially innermost portion of the airfoil.
 10. The airfoil of claim1, wherein each heat pipe comprises a condenser portion, and wherein thecondenser portion of at least one heat pipe of the at least two heatpipes is positioned near a radial midsection of the airfoil.
 11. Theairfoil of claim 1, wherein each heat pipe comprises a condenserportion, and wherein the condenser portion of at least one heat pipeextends into a cavity defined by an outer band of the airfoil.
 12. Theairfoil of claim 1, wherein each heat pipe comprises a condenserportion, and wherein the condenser portion of at least one heat pipeextends into a cavity defined by an inner band of the airfoil.
 13. Theairfoil of claim 1, wherein the airfoil is formed from a ceramic matrixcomposite material.
 14. A method for forming a ceramic matrix composite(CMC) airfoil for a gas turbine engine, the method comprising: laying upa CMC material to form an airfoil preform assembly, the airfoil preformassembly defining an airfoil shape having opposite pressure and suctionsides extending radially along a span, opposite leading and trailingedges extending radially along the span, the pressure and suction sidesextending axially between the leading and trailing edges, the leadingedge defining a forward end of the airfoil, the trailing edge definingan aft end of the airfoil, and a trailing edge portion defined adjacentthe trailing edge at the aft end; processing the airfoil preformassembly to produce a green state CMC airfoil; defining at least twopockets in the trailing edge portion, the at least two pockets radiallyspaced apart from one another; and inserting a heat pipe into eachpocket.
 15. The method of claim 14, wherein at least one pocket of theat least two pockets is defined such that the at least one pocketextends at an angle with respect to an axial direction.
 16. The methodof claim 15, wherein the airfoil preform assembly defines adjacent theleading edge a cavity for receipt of a flow of cooling fluid, andwherein a condenser portion of the heat pipe inserted into the at leastone pocket is positioned near the cavity.
 17. The method of claim 14,wherein at least one pocket of the at least two pockets is defined suchthat the at least one pocket extends radially within the trailing edgeportion.
 18. The method of claim 14, wherein the airfoil preformassembly defines adjacent the leading edge a cavity for receipt of aflow of cooling fluid, and wherein laying up the CMC material to formthe airfoil preform assembly comprises laying up a filler pack preform,a cavity preform, and a plurality of CMC plies that wrap around thefiller pack preform and the cavity preform.
 19. The method of claim 18,wherein at least one pocket of the at least two pockets is definedthrough the cavity preform and the filler pack preform.
 20. The methodof claim 18, wherein at least one pocket of the at least two pockets isdefined in the filler pack preform and is located adjacent the cavitypreform when the filler pack preform is laid up with the cavity preformand the plurality of CMC plies to form the airfoil preform assembly.